The present invention relates generally to gas turbine engines, and, more specifically, to cooling thereof.
In a typical gas turbine engine, air is pressurized in a multistage axial compressor, mixed with fuel in a combustor and ignited for generating hot combustion gases which flow downstream through several turbine stages which extract energy therefrom. The turbine stages include airfoils in the form of stator vanes which turn and accelerate the combustion gases into rotor blades which extract energy therefrom.
In a typical high pressure turbine, both the vanes and blades are hollow and supplied with a portion of pressurized air from the compressor which is used for cooling the respective airfoils thereof. Various features are provided for maximizing the cooling effectiveness of the compressor air, which in turn maximizes the efficiency of the engine.
Typical airfoil cooling features include serpentine cooling passages for selectively cooling the different portions of the airfoil from its leading edge to its trailing edge. The passages may include various forms of turbulators which enhance forced convection cooling. The cooling air may be discharged from the airfoils from various holes in the pressure or suction sides thereof or along the tip or trailing edge thereof. Air discharged through the airfoil sidewalls passes through inclined film cooling holes which effect a cooling air film over the outside of the airfoil to protect the airfoil against the hot combustion gases.
The airfoils may include discrete impingement baffles which firstly direct the cooling air in impingement against the inner surface of the airfoil for cooling thereof, with the spent air then being discharged from the airfoil through various ones of the discharge holes. Since nozzle vanes are stationary and are mounted between radially outer and inner bands, the impingement baffles may be assembled therein through either band.
In contrast, the turbine rotor blades are fixedly mounted at their radially inner ends by dovetails to the outer perimeter of rotor disk. Impingement inserts therefor may therefore be inserted therein typically only from the radially outer tip end thereof. Since rotor blades typically have varying twist, camber, and taper from root to tip, the ability to assemble impingement baffles therein is correspondingly limited.
Since turbine airfoils are subject to the hot combustion gases, they are typically made of advanced superalloy materials having high temperature, high strength capability for maximizing engine performance. To create the various internal cooling features in these airfoils a casting process is typically used. Casting, however, is limited in its ability to precisely form the internal cooling features, which therefore limits the efficiency thereof. And, the impingement baffles must still be separately manufactured and suitably installed in the individual airfoils.
The baffles must also be secured therein, which is typically accomplished at solely one end thereof for permitting unrestrained differential thermal expansion and contraction movement between the baffle and the airfoil under the varying temperature environment of the engine. Since the rotor blades are subject to considerable centrifugal force during operation, baffles therefor must be adequately secured for withstanding the high G28 forces therefrom.
Accordingly, it is desired to further improve the internal cooling features of a gas turbine engine airfoil such as stator vanes and rotor blades for further increasing cooling effectiveness, along with additional benefits.